The present invention relates to gas turbine engines and, more particularly, to turbine rotors used therein.
Gas turbine engines are commonly used for aircraft propulsion and for powering various devices located in aircraft or in ground vehicles or a variety of ground installations (by generating electrical power or compressed air). Such engines in essence have, beginning from the engine atmospheric air inlet, a compressor arrangement having rotatable compressor rotors therein followed by a combustor arrangement for receiving the generated compressed air to add fuel thereto and ignite same. The remaining air and combustions products are hot gases, characterized by high temperatures and pressures, which exit the engine through a turbine arrangement having rotatable turbine rotors therein and thereafter through the exit nozzle to the atmosphere. In doing so, the turbine rotors are forced to rotate and, because of being mechanically coupled to the compressor rotors, these latter rotors are also forced to rotate as are any devices that are also mechanically coupled to the turbine rotors. In aircraft engine applications, the escaping hot gases also provide thrust for propulsion.
The turbine, in addition to rotors with a sequence of radially extending shaped blades about the periphery thereof, has corresponding sequences of circularly arrayed vanes with each such sequence provided in a stationary vane structure located between successive rotors. These vanes stabilize and direct the flows of hot gases from one rotor to the next in a manner seeking to optimize the work extracted from those gases in rotating the turbine rotors. Higher temperatures and pressures in the hot gases passing through the vanes and rotor blades permits extracting greater energy therefrom in turning the rotor blades to thereby increase engine efficiency.
However, operating at such increased temperatures and pressures requires materials in structures in the turbine subjected to such hotter gases to able to operate at such temperatures and pressures at the operational rotation speeds encountered. This includes the rotor blades maintaining their mechanical resistance to the stresses thereon due to such hot, pressurized impinging gases by resisting distortion, stretching, elongation or other forms of metal creep, and resisting wear and corrosion due to those gases or other substances entering the engine inlet airflow. One class of materials found suited for such blades has been forming them arrayed about the outer periphery of a ring in a unitary structure made of cast, sometimes directionally solidified, nickel-base superalloys. These bladed rings are joined at the interior of the ring in the ring opening to a rotor core disk, or hub, suited for being mounted on a rotatable shaft or tube and are typically formed of forged high strength superalloys, which may again be nickel-base superalloys. The material used must also have good low and high cycle fatigue properties, good corrosion resistance and good ductility at room and elevated temperatures. This joining of the bladed ring to a corresponding rotor core disk is typically provided through diffusion bonding of the one to the other in a vacuum furnace for diffusion bonding. However, such furnaces and their operation are costly. Thus, there is a desire for another method for joining together such turbine rotor components and the rotor so formed.